GB2043792A

GB2043792A - Turbine shrouding

Info

Publication number: GB2043792A
Application number: GB8006766A
Authority: GB
Original assignee: MTU Motoren und Turbinen Union Muenchen GmbH
Current assignee: MTU Aero Engines GmbH
Priority date: 1979-02-28
Filing date: 1980-02-28
Publication date: 1980-10-08
Also published as: JPS6316566B2; GB2043792B; US4439982A; DE2907748A1; DE2907748C2; JPS55117012A; FR2450344A1; FR2450344B1

Description

1 GB 2 043 792 A 1

SPECIFICATION Improvements in and Relating to Turbo Machines

This invention relates to a turbo machine, especially a gas turbine engine, and, more 70 particularly, to the control of the clearances prevailing between the outer, free rotor blade tips or an outer shroud of the turbine rotor blades and an adjacent turbine casing.

Low power gas turbine engines of both the turboshaft and the turbojet construction are often provided with a reverse-flow, annular combustion chamber and an axial-flow turbine. The highly efficient operating cycles sought, with their high specific outputs or high specific thrusts at 80 moderate fuel consumptions, naturally make for small size also of the turbine driving the compressor. The radial clearance between the bladed rotor wheel and the casing therefore significantly affects the output, or the thrust and the efficiency. If such gas turbines are additionally subjected to frequent abrupt changes in load conditions, it will be necessary to minimize the blade tip clearances not only under steady-state operating conditions but also under transient operating,conditions when the output level is changed. Further as a result of the design of the compressor turbine with its relatively low hub ratio in conjunction with a normally high ratio of the hub bore diameter to the rim diameter of the 95 rotor disc, the thermal expansion of the rotor blade often exceeds one-third of the total expansion of the rotor. Considering, however, that the thermal expansion of the rotor blades-much like the thermal expansion of the nozzle vanes and 100 the casing-will rapidly follow the variations in working gas temperature while the thermal expansion of the rotor disc clearly lags behind, it follows that conventional constructions designed to minimize blade tip clearances and maintain them constant, are not entirely satisfactory. With the nozzle guide vane support arrangement, e.g. it will be practically impossible for the turbine stator to have sufficient thermal mass for engines of said type.

The present invention aims to minimize and maintain the blade tip clearances in axial-flow turbines for turbomachines, especially for gas turbine engines, over as wide an operating range as possible, even under transient operating 115 conditions.

In accordance with the present invention we propose a turbomachine, particularly a gas turbine engine, in which the clearances prevailing between the outer, free turbine blade tips of 120 turbine rotor blades or an outer shroud of turbine rotor blades and an adjacent turbine casing are controlled by means of an annular, sleeve-shaped component and which is capable of radially elastic deformation and is cooled by air tapped at the compressor exit, and a wear ring attached to the sleeve-like component and divided into segments which are arranged such that, under all operating conditions a predetermined clearance exists between adjacent segments, wherein the frontal area of the segmented wear ring and its immediate means of suspension from the elastic, sleeve-shaped component are adapted to attune the heating rate to the thermal expansion of the turbine rotor disc so ensuring thermally elastic expansion of the sleeve-shapedcomponent; and wherein, the material and configuration of the turbine casing or of the nozzle vane support are attuned, especially with regard to the arrangement of the thermal insulation liners, to the size and the rate of thermal expansion of the turbine rotor disc.

Under ail operating conditions the segmented wear ring exhibits a certain clearance circumferentially between its various segments. A certain radial clearance likewise exists between the blade tips and the wear ring. At rising engine speed-where the rate of change in input is for the present considered negligible-the turbine casing or nozzle guide vane support and the elastic sleeve enlarges radially as a function of the compressor exit temperature, while the ring segments enlarge circumferentially as a function of turbine entry temperature. At the same time, heat will flow from the segmented wear ring, over its frontal area, into the elastic sleeve to heat the suspension area of the wear ring segments and elastically expand this component. The gap prevailing between the rotor and the wear ring, then, is ultimately controlled on the casing side by thermal expansion and elastic deformation of the elastic sleeve in combination. When the output of the engine is reduced, the thermal expansion profile of the segmented wear ring is reversed analogously.

Abrupt load variations,e.g. when accelerating from a lower output level to a higher one, cause the segments of the wear rings, as well as the turbine rotor blades, very rapidly to follow the inertia temperature associated with the new steady-state operating condition and expand circumferentially. Owing to the clearance between the segments circumferentially the diameter of the segmented wear ring is controlled by the thermal expansion of the sleeve in the segment suspension area, which is supplied with air bled from the compressor exit. Owing to its thermal mass and perhaps to thermal insulation provisions the thermal expansion of the nozzle vane support occurs at a relatively large time constant. Considering that the compressor exit temperature spontaneously follows the new steady-state load point of the engine, this thermal expansion takes place relatively slowly, i.e. about as rapidly as the thermal expansion of the rotor blades. The heating and the attendant thermal expansion of the elastic sleeve as heat is transmitted from the wear ring segments, occurs after a certain delay corresponding to the delay with which the rotor disc will heat up after the rotor blades. The rate at which this occurs is controlled by suitably selecting the size of the frontal area of the segmented wear ring and by the configuration of the mating areas between the 2 GB 2 043 792 A 2 segments and the elastic sleeve for the respective engine application.

When the load is reduced abruptly, the various processes will be reversed: The segments of the wear ring first follow, at approximately the same rate as the rotor blades, the inertia temperature corresponding to the new operating condition, which casues them to shrink circumferentially. Simultaneously the cooling of the elastic sleeve with air from the compressor exit causes radial shrinkage in the segment suspension area, which in turn causes a reduction in the diameter of the segmented wear ring at a faster rate, i.e. similar to the rate of the shrinkage of the rotor blades.

Thermal balance of the segment suspension area will then occur and cause further cooling of the elastic sleeve and thus, a reduction in the diameter of the segmented wear ring, yet at a slower rate, i.e. as rapidly as the shrinkage of the rotor disc. This configuration of the turbine stator thus ensures that the gas turbine engine can be operated at narrow rotor tip clearances over a wide range and also under transient operating conditions. which automatically improves the performance and the efficiency as compared with conventional engines.

An embodiment of the invention will now be described by way of example with reference to the accompanying drawings in which:

Fig. 1 is a part longitudinal section of a gas turbine engine according to the present invention, and, Fig. 2 is a cross-section taken on line 11-11 of Fig. 1.

The gas generator of the engine of Fig. 1 100 comprises a centrifugal compressor 1 followed by a radial-flow diffusor 2 from which the flow issuing from the compressor V is deflected axially through approximately 90-degrees to an axial- flow cascade 4 downstream of the bend 3. Compressor air V from the axial- flow cascade 4 enters a first annular duct 5 and, after flowing around the combustion chamber head, reaches a second annular duct 6, both ducts being coaxial with the longitudinal centerline 7 of the engine. The first and second ducts 5 and 6 are defined on the one hand by the flame tube walls 8 of an annular, reverse flow combustion chamber 9 coaxial with the longitudinal centerline 7 and on the other hand by an outer casing wall 10 and a nozzle guide vane support 11 formed in continuation of and connected to the casing wall 10.

The vaulted rear wall of the flame tube, which is not shown on the drawing, is surrounded at a distance by the casing wall 10, which extends parallel to it. A portion of the incoming compressor air V is admitted for combustion through several, circumferentially equally spaced vaporizer tubes, exemplified by the numeral 12, which are connected to the rear wall of the flame tube.

From the annular duct 6 the compressor air V flows into an annular duct 13 which connects directly to the duct 6 and which radially expands 130 relative to the flame tube walls 8. From here the air is routed to serve various cooling functions in a manner more fully explained elsewhere herein.

The gas generator of the engine further comprises a two-stage turbine driving the compressor, its nozzle guide vanes and rotor blades being indicated serially from left to right by the numerals 14, 15 and 16, 17 respectively. The two-stage compressor turbine also comprises two turbine rotor discs 19, 20 rigidly connected for rotation together by, inter alia, circumferentially arranged teeth 18, with the turbine rotor disc 19 being coupled to the disc 24 of the centrifugal compressor 1 through further rotor components 21, 22 and circumferentially arranged teeth 23.

In Fig. 1 the numeral 25 indicates a tie rod, in the form of a tubular shaft, for the gas generator groups. A tubular shaft 26 is carried through this tie rod to transmit the power output of a mechanically independent turbine arranged downstream of the compressor turbine, to a gearbox forward of the engine.

Also in Fig. 1, an annular, sleeve-shaped component 27 capable of radially elastic deformation is suspended from the nozzle guide vane support 11. This sleeve 27 is supplied with air V from the compressor exit and directed, from annular duct 13 and a perforated sleeve 28, against the sleeve 27 in high-energy jets A (impingement cooling). The perforated sleeve 28 is connected to the sleeve 27 and also to a deflector bend 8' and co-operates with the sleeve 27 to support the turbine inlet vanes 14. The deflector bend 8' is loosely suspended by its upstream end in a forked section G of the flame tube walls 8. The components 81,14, 27 and 28 form a constructional unit. From the annulus 29 enclosed by the perforated sleeve 28 and the sleeve 27 a portion of the incoming compressor air is tapped for cooling (arrowhead K) into the hollow vanes 14 and then returned to the gas stream.

Attached to the sleeve 27 is a wear ring 30 divided into segments 31, 32, 33 (Fig. 2) between the adjacent abutting edges of which, under all operating conditions, exists a certain clearance S where the various segments 31, 32, 33 of the wear ring 30 are circumferentially sealed one with the other by means of connecting plates 34.

With reference now to Fig. 2 the wear ring 30 is segmented in suitably selected sequence has erodable liners 35, 36, 37.

With reference again to Fig. 1 the segmented wear ring 30 is cooled by impingement (arrowheads B), of a remaining portion of the air taken from the compression outlet and flowing into an annulus 38 from the annular duct 13 in the direction of arrow R. This annulus 38 is defined between a further perforated sleeve 39, the sleeve 27, and the ring 40 attached to the nozzle vane support 11. Said support ring 40 supports the rear end of the wear ring 30 and the turbine nozzle vanes 16. Also supported by the ring 40 is the rear end of the further perforated sleeve 39. A residual air stream R' issuing from i w i 3 GB 2 043 792 A 3 the annular duct formed-between the wear ring and the further perforated sleeve 30 serves both a sealing air function and to provide film cooling along that surface of the wear ring 30 which faces the blade tips. Another portion T diverted from the annular duct 13 breaks down into a cooling air stream U for the vanes 16 and into a further sealing or film cooling air stream W for the second stage of the turbine.

As it will become apparent from Fig. 1 the segmented wear ring 30 is suspended by axially projecting ends 41 from colar-shaped steps 42 of the sleeve 27.

The cross-hatched frontal areas of the 55 segmented wear ring 30 and its immediate suspension means (ends 41) on the elastic sleeve 27 have dimensions selected to match the heating profile with the thermal expansion of the turbine rotor disc 19 to ensure thermally elastic expansion of the sleeve 27; and the material and configuration of the turbine nozzle vane support 11 is selected, especially with regard to the disposition of thermal insulation liners, to suit the amount and rate of thermal expansion of the turbine rotor disc 19.

The turbine nozzle vane support 11 (Fig. 1) is also supp[led and cooled with air V taken from the compressor outlet, a layer Js of thermal insulation such as a ceramic material being arranged for best results on that side of the vane support 11 facing away from the stream of compressor air V. The segmented wear ring 30 can likewise be fitted with similar layers of thermal insulation Js', on the surface confronting the rotor blade tips.

These layers of thermal insulation Js, Js' act to delay thermal effects arising from the extraction of air taken from the compressor outlet.

Claims

1. A turbomachine in which the clearances prevailing between the outer, free turbine blade tips of turbine rotor blades or an outer shroud of turbine rotor blades and an adjacent turbine casing are controlled by means of an annular, sleeve-shaped component suspended from the outer turbine casing or an associated nozzle vane support and which is capable of radially elastic deformation and is cooled by air tapped at the compressor exit, and a wear ring attached to the sleeve-like component and divided into segments which are arranged such that, under all operating conditions a predetermined clearance exists between adjacent segments, wherein the frontal area of the segmented wear ring and its immediate means of suspension from the elastic, sleeve-shaped component are adapted to attune the heating rate to the thermal expansion of the turbine rotor disc so ensuring thermally elastic expansion of the sleeve-shaped component; and wherein, the material and configuration of the turbine casing or of the nozzle vane support are attuned, especially with regard to the arrangement of the thermal insulation liners, to the size and the rate of thermal expansion of the turbine rotor disc. 65

2. A turbomachine according to Claim 1, wherein the turbine casing or the nozzle vane support is also supplied with air bled from the compressor exit.

3. A turbomachine according to Claim 1 and 2, wherein the nozzle vane support or the turbine casing and/or the segmented wear ring are thermally insulated to delay the thermal effect exercised by the air from the compressor exit.

4. A turbomachine according to any one of Claims 1 to 3, wherein the nozzle vane support forms part of the combustion chamber outer casing of an annular, reverse-flow combustion chamber arranged coaxially with the engine centerline to surround the turbine.

5. A turbomachine according to any one of Claims 1 to 4, wherein the sleeve-shaped component and/or the segmented wear ring are cooled by the impingement of air.

Printed for Her Majesty's Stationery Office by the Courier Press, Leamington Spa, 1980. Published by the Patent Office, 25 Southampton Buildings, London, WC2A 1 AY, from which copies may be obtained.